(1) Field of the Invention
The present invention relates to skid landing gear having at least one damper, and also to an aircraft including the landing gear. More precisely, the invention lies in the technical field of rotorcraft skid landing gear.
(2) Description of Related Art
Conventionally, a rotorcraft has landing gear by which the rotorcraft stands on the ground. More particularly, such landing gear includes “skid” landing gear having first and second longitudinal support skids. The skids are for coming into contact with the ground and they are arranged on either side of the fuselage of the rotorcraft. The first and second longitudinal support skids together define a plane referred for convenience as the “support plane”.
The aircraft then stands on the ground via two elongate skids.
In order to connect each skid to the fuselage of the aircraft, the skid landing gear may be provided with first and second transverse cross-members, each interconnecting the first and second skids.
The first cross-member is said to be the “front” cross-member since the first cross-member connects the fuselage to zones that are situated towards the front of the first and second longitudinal skids. Conversely, the second cross-member is said to be the “rear” cross-member insofar as the second cross-member connects the fuselage to zones situated towards the rear ends of the first and second longitudinal skids.
The landing gear is then fastened to the aircraft via the front and rear cross-members. A cross-member is a frame of structure that extends transversely.
This kind of landing gear enables rotorcraft to stand on surfaces of a wide variety of types.
Furthermore, rotorcraft landing gear may be mainly subjected to two types of stresses during landing: vertical stress associated with vertical forces and movements that are directed in a vertical direction, and roll and pitching stresses that are associated with roll and pitching forces resulting from roll and pitching movements of the rotorcraft.
These stresses, in particular in roll and in pitching, can give rise to the ground resonance phenomenon on a rotorcraft that has a main rotor with hinged blades for providing it with lift.
Under certain very particular conditions, lead/lag resonance modes of the blades can become coupled in unstable manner with the movement of the rotorcraft fuselage in elastic deformation modes of the landing gear, in particular in roll: this gives rise to the phenomenon known to the person skilled in the art as “ground resonance”.
During the ground resonance phenomenon, the blades are subjected to lead/lag movement, i.e. to movement in the plane of the rotor. This movement is out of phase among the various blades and therefore has the consequence of generating unbalance by shifting the center of gravity of the rotor away from the axis of rotation of the rotor.
This unbalance has the effect of causing a rotorcraft standing on the ground on its landing gear to be subjected to excitation (rotating force). This excitation causes the center of the rotor to move in the plane of the rotor and thus causes the fuselage to move. Under such circumstances, the movement of the center of the rotor once more excites lead/lag movement of the blades. The fuselage-and-rotor assembly then becomes a system that is fully coupled and divergent. Under certain conditions, the rotorcraft can overturn in only a few seconds in the absence of appropriate means or decisions.
Concerning these conditions, the rotor, when isolated from the fuselage, has its own lead/lag resonant frequency ωδ. In a stationary reference frame and for the rotor rotating at a frequency Ω, the excitation frequency due to the lead/lag movement of the blades (oscillations at their resonant frequency ωδ) is equal to |Ω±ωδ|.
Under such circumstances, roll instability may occur in particular if the resonant frequency of the fuselage on its landing gear is close to |Ω−ωδ|.
In particular, the resonant frequency of a mode of vibration of the fuselage crossing the lead/lag resonant frequency of the blades generates coupling that is unstable if the damping of said mode of vibration of the fuselage, when coupled to the lead/lag mode of the blades, is negative.
In order to avoid instability, a manufacturer may seek to avoid these frequencies crossing or to obtain such a frequency crossing at a speed of rotation of the rotor that does not run any risk of leading to instability. In order to obtain this result, the manufacturer can adapt the stiffness in roll and/or in pitching of the landing gear.
Nevertheless, it can be difficult to adapt landing gear. In particular, a compromise needs to be found between firstly the vertical stiffness of the landing gear that governs comfort and the level of load imparted to the structure during a landing, and secondly the stiffnesses in pitching and in roll, which have a large influence on ground resonance behavior.
The person skilled in the art uses the term “vertical stiffness” to designate the stiffness of the landing gear under the effect of any vertical stress along the axis in elevation of the aircraft, and assuming the rotorcraft has an attitude that is static.
Skid landing gear is thus generally lengthy and difficult to develop. This development is therefore rarely put into question during the lifetime of the aircraft.
Nevertheless, substantial modifications to an aircraft may take place during its lifetime, e.g. leading to an increase in the weight of the aircraft. The resonant frequencies of the fuselage in roll and/or in pitching can then vary, thereby running the risk of the ground resonance phenomenon appearing.
Under such circumstances, a manufacturer may be tempted to modify the roll and/or pitching stiffnesses of the landing gear without having too much influence on the behavior of the aircraft, in particular while landing.
For this purpose, geometrical modifications may be made to a skid landing gear. Nevertheless, such geometrical modifications may suffer the drawback of modifying the vertical stiffness of the landing gear. The behavior of the landing gear may then in particular be impacted.
In order to avoid the ground resonance phenomenon then appearing, a manufacturer may seek to damp the vibration modes of the fuselage, specifically for the purpose of pushing back the stability limit to speeds of rotation of the rotor that cannot be reached while on the ground. This improvement in damping is also advantageous since it provides an improvement in robustness, but such positioning of the vibration mode frequency can sometimes be difficult to optimize for all of the configurations of the aircraft and for all types of ground landing that are possible for a rotorcraft (in particular landing on a slope).
For example, a damper is known that is arranged directly between a skid and a fuselage. The effectiveness of that device is moderate given the relatively small movement between two attachment points.
It is nevertheless difficult to add damping because of the small amplitudes of the movements of the landing gear relative to the fuselage. In addition, the arrangement of the damper can increase the loads exerted on the structure of the fuselage connected to the damper, and that can raise a problem when installing a damper on a structure that has not been designed for that purpose.
Document U.S. Pat. No. 6,244,538 describes landing gear.
That document makes it possible to position the resonant frequencies in roll and in pitching of the fuselage relative to the excitation frequencies as a function of the center of rotation defined by the support links.
Document FR 2 554 210 describes a flexible beam made of composite materials that is substantially in the form of an elongate box with a laminated structure. Two rigid soleplates are connected together by two webs.
A deformable energy-absorbing pad is arranged between the two soleplates and includes at least one block of elastomer material having high remanence in deformation.
The beam also has at least one viscoelastic damper mounted on the outside face of the soleplate. That damper is stressed in traction by links during bending deformation of the beam so as to produce damping that is additional to the damping provided by each energy absorber pad.
That configuration suggests using an elastomer within each cross-member and within each damper fastened under the cross-member.
Document U.S. Pat. No. 4,270,711 describes landing gear provided with a beam connected by a pivot to a cross-member of the landing gear in such a manner as to be capable of pivoting about an axis. The ends of the beam are then fastened to the structure of the aircraft.
Document U.S. Pat. No. 3,173,632 describes landing gear having two skids that are connected together by two torsion rods. Each torsion rod is secured to two arms, each arm being hinged to an upright bracket extending in elevation from a skid.
Furthermore, movement-preventing means can allow or prevent each torsion rod from pivoting about its axis of symmetry.
Document FR 2 995 874 describes aircraft landing gear having first and second skids together with two cross-members. Each cross-member comprises a first branch secured to the first skid and a second branch secured to the second skid, together with a central portion that is secured to the first and second branches, which are branches that extend downwards. The landing gear has at least one stiffener with at least one link and at least one means for limiting deformation in roll of the central portion of a cross-member, each limitation means being secured to said central portion of the cross-member, at least one main hinge hinging each link to said limitation means and a secondary hinge hinging each link to a point that is outside the central portion in order to limit the deformation of the central portion as a result of roll movement of an aircraft.
Landing gear is also known that has first and second skids, a front cross-member, and a rear cross-member. That landing gear includes at least one stiffener arranged on a cross-member, said stiffener having two rockers, each rocker having an outer end secured to the cross-member. The landing gear includes two hinge means for hinging each rocker to a carrier structure and an elongate link member extending from a first end of the hinged connection to the first rocker to a second end of the hinged connection to the second rocker.
Those rockers therefore seek to stiffen the cross-member.
Also known is Document U.S. Pat. No. 4,519,559.
Documents EP 0 512 898, U.S. Pat. No. 3,822,048, and U.S. Pat. No. 3,716,208 are also known.